Efforts to develop the next generation of aircraft with ever-increasing levels of performance – higher, farther, faster, cheaper – face great technical challenges. One of these technical challenges is to reduce structural weight of the aircraft. Another is to look to aircraft configurations that have been unrealizable to date. Both of these paths can lead to a rigid flex coupling phenomenon that can result in anything from poor flying qualities to the loss of an aircraft due to flutter. This has led to a need to develop an integrated flight and aeroelastic control capability where structural dynamics are included in the synthesis of flight control laws. Studies have indicated that the application of an integrated flight and aeroelastic control approach to a SensorCraft high-altitude long-endurance vehicle would provide substantial performance improvement(1,2). Better flying qualities and an expanded flight envelope through multi-flutter mode control are two areas of improvement afforded by integrated flight and aeroelastic control. By itself, multi-flutter mode control transforms the flutter barrier from a point of catastrophic structural failure to a benign region of flight. This paper discusses the history and issues associated with the development of such an integrated flight and aeroelastic control system for the X-56A aircraft.

Extensions to Maskell’s original correction method, developed over several years, are consolidated and designated ‘Maskell III’. The procedures were applied to dedicated tests on a family of flat-plate wing models in a small, low-speed wind tunnel at NRC. Test conditions included angles of attack from -10° to 110° and models of up to 16% of tunnel area. Off-centre tests were included with model-to-wall distances down to 0.72 chords. Corrected lift and drag data correlated well between models of different sizes and at different offsets from the tunnel centreline. Comparisons are made with corrections using the pressure-signature and two-variable methods, emphasising post-stall conditions. These showed that the ‘Maskell III’ procedures, which require minimal input, correlated as well as the other methods for most model sizes and positions in the tunnel.

A lecture called “Some aspects of buffeting” was given to the Society in 1961 by my colleague Mr. R. Fail. In that lecture the emphasis was primarily on the prediction of buffet onset in flight from model tests in wind tunnels, and on the prevention of buffeting. In the subsequent 11 years there has been much more interest in the prediction of the severity of buffeting above the buffet boundary, for both transport and fighter type aircraft, if buffeting cannot be prevented or alleviated. I have tried to reflect this interest and I hope that the title of this paper will also suggest a cautious examination of the physical processes at work above the buffet boundary, when the boundary layer has separated. We still really know very little about these processes but I hope that the paper may stimulate further research and questioning.

Measurements of momentum in regions of turbulent mixing show that an analogy exists between the processes of turbulent and molecular balance which enables a rational application of the differential equation of thermal conduction to the propagation of turbulent momentum to be made. This fact forms the basis of a new phenomenological theory of free turbulent flows.

Computational fluid dynamics models range from the finite difference type grid-based method to the Lagrangian style vortex cloud simulation technique for solving the Navier-Stokes equations. This paper undertakes a comparison of these two methods for the classical datum bluff body case of flow past a stationary circular cylinder at low Reynolds numbers in the range 10 to 220. Comparisons include time-history, time-mean and root-mean-square values of oscillating drag and lift coefficients, frequency of vortex shedding and related vortex street wake flow patterns. Particularly close agreement was obtained for Strouhal number versus Reynolds number, and good agreement for time-mean value of drag coefficients; comparison was also made with experimental results. Attempts are also made to calculate the skin friction and surface pressure components of the cylinder drag, revealing the significance of skin friction drag within this range and its relative insignificance above a Reynolds number of 220.

High temperatures affect buckling because the properties of the materials change and because thermal stresses and creep develop. A survey is given of the known solutions of problems arising in consequence of these phenomena and new theories of the creep buckling of columns and of thin circular cylindrical shells are presented.

Transverse curvature effects in axi-symmetric flow on a circulai cylinder are discussed and it is deduced that for small δ/a, where δ is the boundary layer thickness and a is the cylinder radius, transverse curvature effects are small, for δ/a = 0(1) the effect is felt mainly in the outer region and for δ/a » 1, both the inner and the outer regions are affected. It is argued that the last case can only be achieved experimentally in the laboratory if the radius Reynolds number Ra= U1a/v is small and the streamwise Reynolds number Rx=U1xlv is large, implying x/a large. New experimental data are presented for this case. The data show that both the mean and turbulent flow field scale on outer variables throughout the layer and, for Ra≥455, exhibit Reynolds number independence. No logarithmic region is found in the mean profiles, a result which is echoed by the absence of a shoulder in the streamwise turbulence profiles which clearly evidence that the flow in the layer is fully turbulent. A flow model is suggested, in physical terms, to explain the very high turbulence levels found close to the cylinder at values of Ra which should, according to stability theory, be bordering on laminar flow.

The development of high-performance computing and computational fluid dynamics methods have evolved to the point where it is possible to simulate complete helicopter configurations with good accuracy. Computational fluid dynamics methods have also been applied to problems such as rotor/fuselage and main/tail rotor interactions, performance studies in hover and forward flight, rotor design, and so on. The GOAHEAD project is a good example of a coordinated effort to validate computational fluid dynamics for complex helicopter configurations. Nevertheless, current efforts are limited to steady flight and focus mainly on expanding the edges of the flight envelope. The present work tackles the problem of simulating manoeuvring flight in a computational fluid dynamics environment by integrating a moving grid method and the helicopter flight mechanics solver with computational fluid dynamics. After a discussion of previous works carried out on the subject and a description of the methods used, validation of the computational fluid dynamics for ship airwake flow and rotorcraft flight at low advance ratio are presented. Finally, the results obtained for manoeuvring flight cases are presented and discussed.

A study of shock tube operation with free-piston compression of the driver gas is presented. It is shown theoretically and experimentally that shock Mach numbers exceeding 30 in air may be obtained with an apparatus of moderate size. The operating limitations imposed by the maximum pressures allowed in the apparatus, by the ratio of the volume of the shock tube to the volume of the compression tube, and by the maximum pressure available for driving the piston are all explored theoretically. The compression process is studied experimentally, and results for three different compression tubes are used to indicate empirical rules for maximising the efficiency of the compression process. Streak photographs of the movement of the luminous gas behind the shock wave in a 1 in diameter shock tube are used to demonstrate the quality of the shock-wave motion produced.

Convective heat transfer data is frequently obtained from transient surface temperature measurements. Thin film resistance gauges, thermocouples, and thermochromic liquid crystals, are used in various situations to measure the surface temperature history. By assuming that uniform semi-infinite flat plate conditions apply, it is possible to express the instantaneous surface heat flux as an analytical function of the transient surface temperature(1). Various approaches can be used to account for the presence of multi-layered substrates and finite thickness substrate effects (Schultz and Jones(1); Doorly and Oldfield(2); Guo et al(3)), however, the effects of surface curvature are usually neglected.

In the aeronautical field, there has been from time to time an interest in inflatable structures, ranging from their consideration for use in space to the design and manufacture of an inflatable wing for a man-powered aircraft and their use in remotely piloted vehicles . This interest stems from the fact that such structures have relatively low weight; are suitable for reacting light loads and can be packed into a small volume.

One particularly useful structural component is the inflated circular cylinder acting as a cantilever and a number of such cylinders are used in Ref. 3 to form the main load carrying structure of a wing. In Ref. 1 an expression is derived for the collapse load of a single cylinder with the assumption that the root behaves like a plastic hinge, and in Ref. 5, Comer and Levy go on to consider the beam deflections and stresses between incipient wrinkling and final collapse.

PIV measurements in the near-field region of a jet flow emanating from a round synthetic jet actuator into quiescent air were conducted over a range of operating conditions. The primary purpose of this work was to investigate the nature of synthetic jets at different operating conditions and to examine the jet flow parameters that dictate the behaviour of synthetic jet actuators. The effects of varying diaphragm displacement and oscillatory frequency for fixed actuator geometry were studied. It was observed that the characteristics of synthetic jets are largely determined by the Reynolds number and stroke length. An increase in the former is observed to increase the strength of consecutive vortex rings that compose a synthetic jet, whereas an increase in the latter results in an increase in relative vortex ring spacing and for further increases in stroke length, shedding of secondary vortices. Correlations were also made between the operating parameters and the performance parameters most effective for flow control and which therefore determine the impact of a synthetic jet on an external flow. Relations of time-averaged dimensionless mass flux, momentum flux and circulation with the jet flow conditions were established and found to widely support an analytical performance prediction model described in this paper. It is anticipated that the experimental data obtained in this study will also contribute towards providing a PIV database for macro-scale synthetic jet actuators.

The actuation mechanism is a crucial aspect in the design of morphing structures due to the very stringent requirements involving actuation torque, consumed power, and allowable size and weight.

In the framework of the CRIAQ MD0-505 project, novel design strategies are investigated to enable morphing of aeronautical structures. This paper deals with the design of a morphing aileron with the main focus on the actuation technology. The morphing aileron consists of segmented 'finger-like' ribs capable of changing the aerofoil camber in order to match target aerodynamic shapes. In this work, lightweight and compact actuation kinematics driven by electromechanical actuators are investigated to actuate the morphing device. An unshafted distributed servo-electromechanical actuation arrangement is employed to realise the transition from the baseline configuration to a set of target aerodynamic shapes by also withstanding the aerodynamics loads. Numerical investigations are detailed to identify the optimal actuation architecture matching as well as the system integratability and structural compactness.

The increasing environmental requirements in the air transport sector pose great challenges to the aviation industry and are key drivers for innovation. Besides various approaches for increasing the efficiency of conventional gas turbine engines, electric propulsion systems have moved into the focus of aviation research. The first electric concepts are already in service in general aviation. This study analyses the potentials of electric and turbo hybrid propulsion systems for commercial aviation. Its purpose is to compare various architectures of electrical powertrains with a conventional turboprop on a regional aircraft, similar to the ATR 72, on engine and flight mission levels. The considered architectures include a turbo-electric (power controlled and direct driven), hybrid-electric (serial and parallel) and a pure electric concept. Their system weights are determined using today’s technology assumptions. With the help of performance models and flight mission calculations the impact on fuel consumption, CO ${}_{2}$ emissions and aircraft performance is evaluated.

To obtain profiles of shear stress and eddy viscosity, the boundary-layer equations in finite-difference form have been applied to published turbulent-boundary-layer developments measured in nominally two-dimensional conditions. In applying this procedure, measured velocity profiles have been represented by members of Thompson’s profile family. Except close to separation, these representations are very satisfactory, and derived shear-stress profiles are generally in good agreement with direct measurements. Various eddy-viscosity and mixing-length models are compared with the results of the analysis and are found in general to differ widely among themselves and from the present results. The widely used assumption, l = ky in the wall region, appears to be invalid.

A theoretical investigation of short-scale inviscid instability in boundary-layer flow past an obstacle on a surface is described, with the application to trip-wire transition and wall roughness effects in mind. An analytical model and computational results show that the occurrence of an inflexion point in a local velocity profile is necessary but not sufficient for the instability to arise.

The unsteady breakdown or stall of streamlined flow near the rounded leading edge of an aerofoil, as the small angle of attack is raised above its critical stall value, is studied theoretically for large Reynolds number motion. The unsteady developments take place first over a relatively slow time scale but then the corresponding solution breaks down with a singularity, forcing a switch to a faster and more nonlinear process. The latter involves a very pronounced local bulge appearing in the flow displacement, accompanied by reversed flow at the aerofoil surface, and comparisons with experimental observations of dynamic stall are noted.

In the coming century, the impact of air travel on the environment will become an increasingly powerful influence on aircraft design. Unless the impact per passenger kilometre can be reduced substantially relative to today’s levels, environmental factors will increasingly limit the expansion of air travel and the social benefit that it brings. This essay considers the three main impacts, noise, air pollution around airports and influence on climate change. Of the three, impact on climate change is taken to have the greatest long-term importance and is discussed at the greatest length. It is argued that, of the three main contributors to climate change from aircraft – CO2 emissions, NOX emissions and the creation of persistent contrails – it is the last two which are the most promising targets. Ways of reducing the impacts of these two are discussed and it is noted that, in each case, the best environmental result is likely to entail some increase in CO2 emissions. It follows that regulatory or economic measures to reduce impact on climate should be framed so as to do just that. Measures framed purely in terms of CO2 emissions are likely to be counter-productive. Nevertheless, the design of aircraft to reduce fuel burn and hence CO2 emission remains a key long-term objective; the essay considers the potential offered by new technology and new design concepts in this arena.

An exact integral formulation of the two-dimensional full potential equation is presented. The well developed standard surface panel method based on the linear Prandtl-Glauert equation is extended by means of sources in the field surrounding the configuration, allowing the calculation of transonic flows.

A major attraction of the Field Panel Method is that solutions can be obtained using standard surface panelling together with a field grid which has no special form. The surface panels are used to satisfy the boundary conditions so that the field grid does not have to be surface-fitted which is a requirement for existing field methods. Further, the grid need only be in a region close to the configuration where nonlinear compressibility effects are non-negligible and need not extend to the far field as generally required. Grid generation is therefore trivial allowing solutions for arbitrary configurations. Convergence, especially for flows with shock waves, is enhanced by use of a modified Approximate Factorisation scheme. Results for single aerofoils are compared with Garabedian-Korn solutions and show good agreement; and to demonstrate the ease of grid generation, results for multicomponent aerofoils are also presented. Finally, an outline is given of how the method can be extended to threedimensions.

The work was carried out in conjunction with the Department of Applied Mathematics, University of Leeds together with Procurement Executive, Ministry of Defence.